Low hub-to-tip ratio fan for a turbofan gas turbine engine

ABSTRACT

A fan for a turbofan gas turbine engine, the fan comprising a rotor hub and a plurality of radially extending fan blades integral with the hub to form an integrally bladed rotor. Each fan blade defines a leading edge. A hub radius (R HUB ) is the radius of the leading edge at the hub relative to a centerline of the fan. A tip radius (R TIP ) is the radius of the leading edge at a tip of the fan blade relative to the centerline of the fan. The ratio of the hub radius to the tip radius (R HUB /R TIP ) is at least less than 0.29. In a particular embodiment, this ratio is between 0.25 and 0.29. In another particular embodiment, this ratio is less than or equal to 0.25.

TECHNICAL FIELD

The present invention relates to turbofan engines and more particularlyto fans for such engines having low hub to tip ratios.

BACKGROUND

Most gas turbine engine fans are composed of a central hub onto which aplurality of separately formed fan blades are secured. Integrated bladedrotor (IBR) fans are known for their relative lightness and thereforeare desirable, however known IBR fans cannot be formed having a low hubto tip radius ratio because of limitations in manufacturingcapabilities. Such a low hub to tip radius ratio is however desirablebecause it means the maximum diameter of the fan can be reduced withoutnegatively effecting performance. Reducing the overall diameter of thefan reduces weight and improves the efficiency of the fan.

Therefore, while the advantages of reducing the ratio of the radius ofthe hub to the radius of the tip are well appreciated in terms ofreducing the specific flow of air entering the leading edge of the fan,attempts to date to reduce the specific flow by reducing this ratio havenot been readily possible, particularly for IBR fans. Attempts tomanufacture an integrated bladed rotor (IBR) fan with a low hub to tipratio have not been successful because of the lack of space for machinetools between the roots of the blades when the hub is also reduced insize.

SUMMARY

There is therefore provided a fan for a turbofan gas turbine engine, thefan comprising a rotor hub and a plurality of radially extending fanblades integral with the hub to form an integrally bladed rotor, eachfan blade having a leading edge, a hub radius (R_(HUB)) which is theradius of the leading edge at the hub relative to a centerline of thefan, and a tip radius (R_(TIP)) which is the radius of the leading edgeat a tip of the fan blade relative to the centerline of the fan, andwherein the ratio of the hub radius to the tip radius (R_(HUB)/R_(TIP))is at least less than 0.29.

In a particular embodiment, the ratio R_(HUB)/R_(TIP) is less than orequal to 0.25.

In another particular embodiment, the ratio R_(HUB)/R_(TIP) is between0.25 and 0.29.

There is also provided a method of manufacturing an integrally bladedrotor fan for a turbofan gas turbine engine, comprising: forming a rotorhub preform defining a hub radius and having at least a number of rootstubs radially spaced apart on a periphery of the rotor hub perform;providing blade airfoils having a length such that a ratio of the hubradius to a tip radius of the blade airfoils, once mounted to the hub,is at least less than 0.29; and subsequently fastening the bladeairfoils to the root stubs to form fan blades integrally formed with thehub resulting in an integrally bladed rotor fan having a hub to tipradius ratio of at least less than 0.29.

There is further provided a turbofan gas turbine engine comprising a fanupstream of at least one compressor, the fan having a rotor hub and aplurality of substantially radially extending fan blades integral withthe rotor hub to form an integrated bladed rotor, each said fan bladehaving an airfoil defining a leading edge and defining a tip radius(R_(TIP)) which is the radius of a tip of the fan blade at the leadingedge, the rotor hub defining a hub radius (R_(HUB)) which is the radiusof the hub at the blade leading edge, and wherein a ratio of the hubradius to the tip radius (R_(HUB)/R_(TIP)) is at least less than 0.29.

BRIEF DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic cross-sectional view of a gas turbine enginehaving a fan in accordance with the present disclosure; and

FIG. 2 is a partial axial cross-sectional view of an embodiment of thefan of the present disclosure.

DETAILED DESCRIPTION

FIG. 1 illustrates a turbofan gas turbine engine 10 generally comprisingin serial flow communication, a fan assembly 12 through which ambientair is propelled, and a core 13 including a compressor section 14 forpressurizing the air, a combustor 16 in which the compressed air ismixed with fuel and ignited for generating an annular stream of hotcombustion gases, and a turbine section 18 for extracting energy fromthe combustion gases. A centerline main engine axis 13 extendslongitudinally through the turbofan engine 10.

The fan 12 propels air through both the engine core 13 and the bypassduct 22, and may be mounted to the low pressure main engine shaft 11.The fan 12 includes a plurality of radially extending fan blades 20 anda central hub as will be seen, which hub has a nose cone 22 mountedthereto to protect the hub. As will be described in greater detailbelow, the fan 12 is an integrally bladed rotor (IBR), wherein the fanblades 20 are integrally formed with the central hub that is fastened tothe low pressure (LP) engine shaft 11 for rotation therewith.

Referring now to FIG. 2, the IBR fan 12 comprises a plurality of fanblades 20 integrally formed with, and substantially radially extendingfrom, a central fan hub 36 which is mounted to an engine shaft, such asthe low pressure shaft 11, by means of one or more hub support portions38 which are also integrally formed with the hub. Each of the blades 20defines an airfoil 28 which has a leading edge 34 which extends from ablade root 30 to a blade tip 40. The blade 20 is integrated with the hub36, i.e. such the blades 20 are integrally formed as a monolithiccomponent with the fan hub 36 to form an IBR fan. The nose cone 22 ofthe engine may be fastened to an upstream end of the fan hub 36 by aplurality of fasteners 29.

When the radius of the leading edge 30 on the hub 36 is reduced whilethe radius of the blade tip at 40 is maintained, the flow area (FA) ofthe fan 20 is increased thus reducing the specific flow (SF). As seen inFIG. 2, the gaspath through the fan 12 is defined by the annular areabetween the hubs 30 and the tips 40 of the fan blades 20. The radius ofthe fan hub (R_(HUB)), measured at the leading edge 34 of the blade 20,defines the radially inner gaspath boundary and the radius of the bladetip (R_(TIP)), also measured at the leading edge 34, defines theradially outer gaspath boundary. The specific flow of the fan 12 istherefore defined as the mass flow (MF) of air entering the leading edgeof the fan 12, divided by the flow area (FA) at the fan leading edge,normal to the engine axis 13.

The hub to tip ratio of the IBR fan 12 is defined as the ratio of theradius of the fan hub (R_(HUB)) at the leading edge divided by theradius of fan blade tip (R_(TIP)) at the leading edge. As shown in FIG.2, these radii are is measured from the engine centerline axis 13.

Thus, specific flow is determined as follows:

SF=MF/FA,

where SF is the specific flow, MF is the mass flow, and FA the flowarea. Reduction of this SF of the fan is desirable as a reduced SF helpsto improve the overall aerodynamic efficiency of the fan because of thelower air velocity.

A reduction in the hub to tip ratio (R_(HUB)/R_(TIP)) will thereforealso cause a reduction in the specific flow (SF) of the fan.Alternatively, the radius of both the hub 36 and the blade tip 40 can bereduced while retaining the same specific flow SF. However, the ratio ofthe hub to tip radii is preferably reduced. Accordingly, the present IBRfan 12 has a ratio of the hub radius to the tip radius, i.e.R_(HUB)/R_(TIP), which is at least less than 0.29. In a particularembodiment, the ratio of the hub radius to the tip radius(R_(HUB)/R_(TIP)) is between about 0.25 and about 0.29. In a furtherparticular embodiment, the ratio of the hub radius to the tip radius(R_(HUB)/R_(TIP)) is less than or equal to 0.25.

The advantage of a lower tip radius is a smaller diameter fan andtherefore a lighter weight engine. Lowering the hub leading edge radiusalso changes the flow angle of the airstream, and the resulting rearwardsweep in the lower portion of the fan blade airfoils 28 improvesperformance by reducing the leading edge velocities through the sweepeffect and also draws flow towards the hub 36 which helps to reduce flowseparation that the blade root.

The advantage of using the integrally bladed rotor (IBR) fan 12 is itsreduced weight compared to a traditional detachable bladed rotor. Themachining of an IBR fan 12 with such a low hub/tip ratio is madedifficult by the lack of space between the blades 20, particularly atthe blade roots 30 since the gap between the blades is much narrower thesmaller the radius of the fan.

However, in one particular method of manufacturing the IBR fan 12described herein, it has been found that by first machining a root stub44 on the hub 36, the lower hub radius, and more particularly the lowhub to tip radius ratios described above, can be obtained because it iseasier to access the radial gap between adjacent blades 20 with machinetools. The blade airfoils 28 may then be fixed to the root stubs 44 bewelded by Linear Friction Welding (LFW), for example, along the jointline 42 as shown on the blade 20 in FIG. 2. It has been contemplatedthat alternative methods may also be used, such as forming a root stub44 only for every alternate blade, while machining the full blade 20between each alternate root stub. This would allow sufficient access formachine tools between two alternate full blades, to machine around thearound the remaining root stub.

Thus, a low-weight fan 12 as described herein is achieve, because of itsintegrated bladed rotor construction, and which provides a hub to tipradius ratio of at least less than 0.29, and more particularly between0.25 and 0.29, and more particularly still a hub to tip radius ratio of0.25 or less.

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentsdescribed while still falling within the scope of the appended claims,which define the present invention. Such modifications will be apparentto those skilled in the art, in light of a review of this disclosure.

1. A fan for a turbofan gas turbine engine, the fan comprising a rotorhub and a plurality of radially extending fan blades integral with thehub to form an integrally bladed rotor, each fan blade having a leadingedge, a hub radius (R_(HUB)) which is the radius of the leading edge atthe hub relative to a centerline of the fan, and a tip radius (R_(TIP))which is the radius of the leading edge at a tip of the fan bladerelative to the centerline of the fan, and wherein the ratio of the hubradius to the tip radius (R_(HUB)/R_(TIP)) is at least less than 0.29.2. The fan as defined in claim 1, wherein the ratio of the hub radius tothe tip radius (R_(HUB)/R_(TIP)) is less than or equal to 0.25.
 3. Thefan as defined in claim 1, wherein the ratio of the hub radius to thetip radius (R_(HUB)/R_(TIP)) is between 0.25 and 0.29.
 4. A fan preformfor a fan as defined in claim 1, wherein the fan perform includes rootstubs disposed on the hub at positions corresponding to at leastalternate ones of said fan blades, the root stubs being first formed onthe hub prior to blades being fastened thereto.
 5. The fan perform asdefined in claim 4 wherein the root stubs formed on the preform haveairfoils welded thereto to provide the integrally bladed rotor.
 6. Thefan perform as defined in claim 5 wherein the airfoils arelinear-friction-welded to the respective root stubs.
 7. The fan preformas defined in claim 4 wherein all of the fan blades are first formed asroot stubs on the hub.
 8. A method of manufacturing an integrally bladedrotor fan for a turbofan gas turbine engine, comprising: forming a rotorhub preform defining a hub radius and having at least a number of rootstubs radially spaced apart on a periphery of the rotor hub perform;providing blade airfoils having a length such that a ratio of the hubradius to a tip radius of the blade airfoils, once mounted to the hub,is at least less than 0.29; and subsequently fastening the bladeairfoils to the root stubs to form fan blades integrally formed with thehub resulting in an integrally bladed rotor fan having a hub to tipradius ratio of at least less than 0.29.
 9. The method as defined inclaim 8, wherein the step of selecting further comprises selecting thelength of the blade airfoils to define a ratio of the hub radius to thetip radius of the blade airfoils that is between 0.25 and 0.29.
 10. Themethod as defined in claim 9, wherein the step of selecting furthercomprises selecting the length of the blade airfoils to define a ratioof the hub radius to the tip radius of the blade airfoils that is lessthan or equal to 0.25.
 11. The method as defined in claim 8, whereincircumferentially alternate ones of said fan blades are formedintegrally with the hub perform without root stubs, leaving alternateroot stubs on the hub perform to provide access for machine toolsbetween the circumferentially alternate ones of said fan blades.
 12. Themethod as defined in claim 8, wherein the step of fastening furthercomprises welding the blade airfoils to the root stubs using LinearFriction Welding.
 13. A turbofan gas turbine engine comprising a fanupstream of at least one compressor, the fan having a rotor hub and aplurality of substantially radially extending fan blades integral withthe rotor hub to form an integrated bladed rotor, each said fan bladehaving an airfoil defining a leading edge and defining a tip radius(R_(TIP)) which is the radius of a tip of the fan blade at the leadingedge, the rotor hub defining a hub radius (R_(HUB)) which is the radiusof the hub at the blade leading edge, and wherein a ratio of the hubradius to the tip radius (R_(HUB)/R_(TIP)) is at least less than 0.29.14. The turbofan gas turbine engine as defined in claim 13, wherein theratio of the hub radius to the tip radius (R_(HUB)/R_(TIP)) is less thanor equal to 0.25.
 15. The turbofan gas turbine engine as defined inclaim 13, wherein the ratio of the hub radius to the tip radius(R_(HUB)/R_(TIP)) is between 0.25 and 0.29.